Mistake proof identification feature for turbine blades

ABSTRACT

An identification feature is used to unmistakably identify internal features present in different generations of turbine blade designs. The identification feature is located on a root portion of the turbine blade and protrudes to provide a visually identifiable feature that is also readable by a coordinate measuring machine, but does not interfere with installation or operation of the turbine blade. The weight of the identification feature is in a specific proportion to the weight of the turbine blade in order to prevent interfere with operation of the turbine blade during high-speed rotation in a gas turbine engine.

BACKGROUND OF THE INVENTION

This invention relates generally to gas turbine engines and moreparticularly to identification of turbine blades having internalfeatures. In gas turbine engines, fuel is combusted in compressed aircreated by a compressor to produce heated gases. The heated gases areused to turn turbine blades, or airfoils, to produce rotational powerfor, among other things, operating the compressor. During operation ofthe gas turbine engine, temperatures inside the combustion chamber canreach 2500° F., resulting in the blades being subject to temperatures inexcess of 1700° F. In order to cool the turbine blades, relativelycooler compressed air that bypasses the combustion chamber, or bleedair, is forced through internal passages of the blades. The passagesinclude pathways or channels having various geometries in order todirect the bleed air throughout the interior of the blade. The bleed airflowing through the passages maintains a temperature gradient throughoutthe entirety of the blade at which the blade can properly function.

For performance or manufacturing reasons, it is sometimes necessary tochange or modify the interior features of a particular blade model.Meanwhile, the exterior of that blade must be maintained the same inorder to meet the design of the specific gas turbine engines in whichthat model of blade is used. Traditionally, a model number thatidentifies the interior features of the turbine blade is cast on theexterior of the turbine blade casting. The cast model numbers produce ashallowly indented number on the surface of the turbine blade. Theshallow numbers do not create any protrusions or cavities that upset thebalance of the blade while it is rotating. Any, even small,disproportion of weight along the length of the turbine blade canproduce vibrations during the high-speed rotations produced in gasturbine engines.

While the cast model numbers are small enough to prevent anyinterference with the operation or installation of the blade, thenumerals are often illegible and confusingly similar. For example, acast “9” may look like a “0.” Thus, a turbine blade having secondgeneration internal features would look identical to a turbine bladehaving first generation internal features, and there would be nopositive way to identify which generation of internal features itpossesses. Therefore, blades could be improperly introduced into theproduction stream where they would receive incorrect finishingprocedures that are not discovered until a later time. It is desirablefor production cost and safety considerations to completely eliminatethe possibility of these mistakes. There is, therefore, a need for aturbine blade having an identification feature that unmistakablyidentifies the internal features of visually identical turbine bladeswithout interfering with the operation of the blade itself.

BRIEF SUMMARY OF THE INVENTION

The present invention is directed towards a positive identificationfeature used to identify internal features of turbine blades. Theinvention comprises a protruding identification that unmistakablyidentifies the internal features of the turbine blade. The protrudingidentification feature is visually identifiable and readable by acoordinate measuring machine. The protruding identification feature islocated on a root portion of the turbine blade so as to preventinterference with installation of the turbine blade. The protrudingidentification feature weighs approximately 0.1% or less of the weightof the turbine blade in order to prevent interference with operation ofthe turbine blade.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a partially cut away perspective view of a gas turbineengine showing a turbine section in which the present invention is used.

FIG. 2 shows a partially exploded perspective view of the turbinesection of FIG. 1 showing a turbine blade and rotor disc assembly.

FIG. 3 shows the root section of a turbine blade having first generationinternal features.

FIG. 4 shows the root section of a turbine blade having secondgeneration internal features in which the present invention is used.

DETAILED DESCRIPTION

FIG. 1 shows a partially cut away perspective view of gas turbine engine10 showing turbine section 12 in which the present invention is used.Gas turbine engine 10 includes turbine section 12, which is positionedbetween combustion chamber 14 and nozzle 16. Casing 18 shrouds turbinesection 12, combustion chamber 14 and nozzle 16. Turbine section 10 is amulti-stage turbine and includes turbine blades 20A, 20B and 20C; rotordiscs 22A, 22B and 22C and turbine shaft 24. Turbine blades 20A, 20B and20C are radially aligned around the periphery of rotor discs 22A, 22Band 22C, respectively. Rotor discs 22A, 22B and 22C are co-axiallyattached to turbine shaft 24, which extends in an axial direction intogas turbine engine 10.

Fuel is combusted in high-pressure air inside combustion chamber 14 inorder to produce heated gases having high density and high pressure. Asthe heated gases exit combustion chamber 14, they enter turbine section12 at a high velocity. The high-density gases impinge on turbine blades20A, 20B and 20C to produce rotational movement of rotor discs 22A, 22Band 22C, which in turn rotate turbine shaft 24. Rotation of turbineshaft 24 produces mechanical power for driving the compressor section ofgas turbine engine 10. The heated gases continue through turbine section12 and are forced through nozzle 16. Nozzle 16 increases the velocity ofthe gases as they exit gas turbine engine 10 in order to produce forwardthrust for propelling an aircraft.

FIG. 2 shows a partially exploded perspective view of cut away turbinesection 12 of FIG. 1 showing the assembly of turbine blades 20A androtor disc 22A. Turbine blades 20A are radially arranged around theouter circumference of rotor disc 22A. Turbine blades 20A include foil26, shroud 28, platform 30 and root 32. Rotor disc 22A includes slots 34aligned along the outer circumference of rotor disc 22A. Slots 34receive roots 32 of turbine blades 20A. Slots 34 include serrations 36,and roots 32 include tangs 38 having a matching profile with that ofserrations 36. In typical embodiments, roots 32 have a “fir tree” or“dove tail” configuration as is known in the art. Roots 32 are insertedinto slots 34 the axial direction so tangs 38 are locked into serrations36. Tangs 38 and serrations 36 secure turbine blade 20A in the radialdirection during rotation of rotor disc 22A and distribute the loadproduced by the centrifugal momentum of rotating turbine blade 20A.Serrations 36 and tangs 38 also allow for thermal expansion of roots 30and rotor disc 22A in the extreme temperatures reached in gas turbineengine 10. Additionally, rivets or other fastening mechanisms are usedto hold turbine blades 20A in the axial direction.

When turbine blades 20A are inserted into rotor disc 22A, shrouds 28align to form a continuous barrier that assists in preventing gasleakage around the tips of the turbine blade. Shrouds 28 also preventvibration and bending of foils 26. In other embodiments, shrouds 28 arenot used and the blade tips of foils 26 are cut to a knife-edge tip.Similarly, platforms 30 align to form a continuous boundary betweenturbine blades 20A and roots 30.

Typically, bleed air used for cooling turbine blades 20A is introducedthrough an opening located on the bottom of root 32, whereby it enterspassages of an interior cooling system. The interior cooling systemincludes various features and passages in which the bleed air flows. Thebleed air travels through the passages on the interior of turbine blade20A and whisks heat away from foil 26. Typically, the heated bleed airexits the interior of turbine blade 20A through one or more smallorifices 40 located on the trailing edge of foil 26 or on the concavesuction side (not shown) of foil 26.

FIG. 3 shows the root section of turbine blade 20A having firstgeneration internal features. For a particular turbine blade design,changes to the interior features may occur mid-production to increaseperformance of the blade. However, the exterior of every generation ofturbine blade 20A is identical to each other, thereby producing aninterchangeable part that will always fit in the gas turbine engines itwas designed for use in.

FIG. 4 shows root section 32 of turbine blade 20A′ having secondgeneration, or post-modification, interior features in which the presentinvention is used. Once a change has been made to the interior design ofthe model of turbine blade comprising turbine blade 20A, identificationfeature 42 is added to root section 32 to produce turbine blade 20A′.Identification feature 42 provides a mistake proof means fordistinguishing turbine blade 20A from 20A′.

Identification feature 42 provides a positive, raised protuberance thatcan be recognized by visual inspection. Identification feature 42 alsoprovides a feature that can be measured with a Coordinate MeasuringMachine (CMM). During manufacture of turbine blade 20A′ the blade isinspected for dimensional tolerances before being sent for additionalmachining procedures. Identification feature 42 provides a positivefeature that can be included in the dimensional tolerance checklist andchecked for with the CMM. This ensures that the turbine blade beinginspected is in fact turbine blade 20A′ and that it will receivemachining procedures intended for blades with second generation internalfeatures.

The location of identification feature 42 is selected to not interferewith the operation of turbine blade 20A′. For example, it is unfeasibleto put an identifying mark on foil portion 26 because that wouldinterfere with impingement of the hot air on foil 26 and would causevibration of foil 26. For similar reasons, it would be unfeasible to putan identifying feature on shroud 28 or platform 30. Also, it isimpracticable to put an identifying feature in the sides of root portion32 because that would interfere with alignment of serrations 36 andtangs 38. Considering these factors, identification feature 42 is placedon front surface 44 of root portion 32. In other embodiments,identification feature 42 is placed on the rear surface of root portion32. In FIG. 4 identification feature 42 is placed on root portion 32off-center of front surface 44. This moves identification feature awayfrom the parting line of the casting of turbine blade 20A′ and allowsthe mold for turbine blade 20A to be adapted for forming turbine blade20A′. In other embodiments, identification feature 42 is centered onfront surface 44 of root portion 32. Placing identification feature 42on root portion 32 also minimizes the vibration effect caused byidentification feature 42 on foil 26.

To further prevent identification feature 42 from interfering withoperation and installation of turbine blade 20A′, identification feature42 is placed in recess 46 located on front surface 44 of root portion32. Recess 46 is pre-formed into the casting of turbine blade 20A′ forweight reduction purposes or other functional purposes. Additionally,recess 46 can be machined into turbine blade 20A′ for the purposes ofreceiving identification feature 42. Thus, in order to minimize theinterference of identification feature 42 on the installation andoperation of turbine blade 20A′, identification feature 42 does notextend beyond the forward most portion of the leading edge of rootportion 32.

During operation of gas turbine engine 10, rotor disc 22A rotates atspeeds of approximately 15000 revolutions per minute (RPM). During thesehigh-speed rotations the tangential velocity of the tips of turbineblade 20A′ can reach speeds up to Mach 2. Thus, placing even a smallamount of mass on turbine blade 20A′ creates a large force that willinterfere with true rotation of rotor disc 22A and foil 26. Thecentrifugal force generated by the mass of identification feature 42 hasthe potential to create vibrations in the rotation of turbine blade20A′. When the centrifugal forces exert stresses beyond the stresslimits of turbine blade 20A′, especially compounded with resonancevibration, catastrophic failure of turbine blade 20A′ will occur.

Using standard mechanics formulas, the size and mass of anidentification feature 42 that will not cause excessive stresses inturbine blade 20A′ can be determined. It has been determined that whenplacing identification feature 42 on root portion 32, an identificationfeature weighing approximately 0.1% or less of the total weight ofturbine blade 20A will not produce excessive stresses in turbine blade20A′. Therefore, in one embodiment, identification feature 42 weighs0.1% of turbine blade 20A′. For example, if turbine blade 20A′ weighs0.84 lbs., identification feature 42 weighs approximately 0.0008 lbs. orless. This prevents excessive stresses in and vibration of turbine blade20A′ during high-speed rotation of rotor disc 22A during operation ofgas turbine engine 10.

The specific shape of identification feature 42 can have variousembodiments. In FIG. 4, identification feature 42 is a vertical rib. Anadditional vertical rib identification feature 42, or a differentlyshaped identification feature 42, can be added to identify eachsubsequent generation of turbine blade 20A. In various embodiments,identification feature 42 can be circular, star shaped or triangular.The size and shape of each identification feature, or the plurality ofidentification features, is limited by being maintained at or belowapproximately 0.1% of the weight of turbine blade 20A to preventperturbation of turbine blade 20A′ during rotation of rotor disc 22A.The size and shape of identification feature is also limited because itmust not interfere with the installation of turbine blade 20A′.

The present invention has been described as applied to turbine bladesused in the turbine section of a gas turbine engine. The protrudingidentification feature can also be used in rotor blades used in thecompressor section of gas turbine engines or in other rotating foils orblades having varying interior features.

Although the present invention has been described with reference topreferred embodiments, workers skilled in the art will recognize thatchanges may be made in form and detail without departing from the spiritand scope of the invention.

1. A turbine blade comprising: a foil portion having internal coolingfeatures; and a root portion comprising: a leading edge surface; atrailing edge surface; a plurality of locking tangs positioned betweenthe leading edge surface and the trailing edge surface; and asingle-character protruding identification feature weighingapproximately equal to or less than 0.1% of a weight of the turbineblade and located on a recessed surface of the root portion of theturbine blade so as to not extend beyond an outermost surface of theroot portion, the protruding identification feature uniquelyrepresenting a generation of the internal cooling features of theturbine blade.
 2. The turbine of claim 1 wherein the single-characteridentification feature is located on the leading edge surface.
 3. Theturbine blade of claim 2 wherein the single-character identificationfeature is a single elongate rib having a major axis extending in alongitudinal direction with respect to the turbine blade.
 4. The turbineblade of claim 3 wherein the vertical rib is offset from a center of theleading edge of the root portion so as to be offset from a castingparting line.
 5. The turbine blade of claim 1 wherein the internalcooling feature comprises trailing edge cooling holes.
 6. A method forproducing a turbine blade having internal cooling features, the methodcomprising: producing a turbine blade comprising: an airfoil portioncomprising: an exterior gas path surface; and an interior coolingfeature having a specific configuration; and a root portion comprising:a leading edge surface; a trailing edge surface; and a plurality oflocking tangs positioned on side surfaces between the leading edgesurface and the trailing edge surface; producing a single-characterexternal identification feature having a geometry uniquely identifyingthe specific configuration of the interior cooling feature, the externalidentification feature positioned on the root portion of the turbineblade; visually identifying the external identification feature; andperforming machining procedures on the turbine blade correlated to theexternal identification feature.
 7. The method of producing a turbineblade of claim 6 wherein the step of visually identifying the externalidentification feature comprises: measuring the external identificationfeature with a coordinate measuring machine verify the specificconfiguration of the interior cooling feature.
 8. The method ofproducing a turbine blade of claim 7 wherein the step of performingmachining procedures further comprises producing features on the turbineblade corresponding to the verified specific configuration of theinterior cooling feature.
 9. The method of producing a turbine blade ofclaim 8 wherein the step of producing features on the turbine bladecorresponding to the specific configuration of the interior coolingfeature further comprises producing cooling holes into a trailing edgeof the airfoil.
 10. The method of producing a turbine blade of claim 6wherein the step of producing the root portion of the turbine bladefurther comprises: producing a recessed leading edge pocket on the rootportion; and positioning the external identification feature in thepocket such that the identification features does not extend beyond theleading edge surface; wherein the identification feature comprises araised bar extending longitudinally across the recessed leading edgepocket at a position offset from a center of the root portion.
 11. Amethod for producing a redesigned turbine blade, the method comprising:designing a turbine blade comprising: a gas path portion having a firstgeneration cooling feature; and a root portion comprising: a leadingedge surface; a trailing edge surface; and a plurality of locking tangspositioned on side surfaces between the leading edge surface and thetrailing edge surface; redesigning the turbine blade such that the gaspath portion includes a second generation cooling feature; and producingthe turbine blade to include an external identification feature on theroot portion to uniquely indicate the second generation cooling feature.12. The method for producing a redesigned turbine blade of claim 11wherein the external identification feature comprises a single-characterfeature.
 13. The method for producing a redesigned turbine blade ofclaim 12 wherein the external identification feature comprises: a raisedprotrubrance located on a recessed surface of the root portion of theturbine blade so as to not extend beyond an outermost surface of theroot portion.
 14. The method for producing a redesigned turbine blade ofclaim 12 wherein the external identification feature comprises: alongitudinal rib offset from a center of the leading edge of the rootportion so as to be positioned away from a casting parting line of theredesigned turbine blade.
 15. The method for producing a redesignedturbine blade of claim 11 wherein the external identification feature isapproximately equal to or less than 0.1% of a weight of the turbineblade.
 16. The method for producing a redesigned turbine blade of claim11 and further comprising identifying the external identificationfeature using a coordinate measuring machine.